Turbine blades

ABSTRACT

In turbine blades used as nozzle guide vanes  2, 12, 22, 31  for turbine stages of an engine there is a problem with respect to high lift vanes in that there may be significant stagnation point migration about the pressure surface  13, 21, 32  of that vane. In such circumstances, appropriate positioning of coolant flow apertures  11, 24, 34  for coolant film cooling of the vane is difficult. The present invention provides on the leading edge a bulge ridge  37  which limits the range of potential movement for the stagnation point  30  whilst retaining improved lift for higher engine performance. The present invention has particular applicability with flows directed to the turbine NGV which have a significantly variable swirl angle and therefore presentation to the vane  31.

The present invention relates to turbine blades and more particularly but not exclusively to turbine blades used as nozzle guide vanes for a high pressure turbine stage of an engine.

High thermal efficiency is dependent upon high turbine engine temperatures which in turn are limited by turbine blade and nozzle guide vane materials. These nozzle guide vanes can be stationery when they are termed “stator blades” or may-be adjustable in terms of the degree of guiding of air flow presented to the high pressure turbine stage of the engine.

Some turbine nozzle guide vanes include film cooling flows which are designed to cool either the pressure or the suction side of the aerofoil surfaces of the guide vane. Film cooling comprises releasing through a plurality of holes appropriately distributed upon the guide vane surface a flow of cooling air such that a “film” of such cooling air passes over the vane surface in order to cool that vane. The path of the air flows depends upon a leading edge stagnation point upon the aerofoil surface. The location of the leading edge stagnation point splits the main stream flow.

Attached drawing FIG. 1 illustrates a typical nozzle guide vane and turbine blade cooling arrangement 1. Thus, it can be seen that the nozzle vanes 2 are arranged whereby high pressure cooling air emitted from apertures flows over the aerofoil surfaces 3 in order to cool that vane 2. Turbine blades 4 are also cooled by flows through such apertures. As indicated above, a leading edge stagnation point spits the air flow presented to the vanes 2 in the direction of arrowhead A either side of the vane 2.

Attached drawing Prior Art 1 illustrates a plan cross-section of a conventional high pressure nozzle guide vane 12. Thus, as can be seen air flow in the direction of arrowhead AA is split about a stagnation point 10 or stagnation line in three dimensions. The position of the stagnation point 10 is sufficiently well understood to enable the coolant apertures 11 described previously with regard to FIG. 1 to ensure appropriate film cooling of the vane 2. Generally, the stagnation point is well understood as a high pressure nozzle guide vane is subject to relatively small variation in the inlet swirl angle of the flow AA. High pressure nozzle guide vanes are generally insensitive to such variations in the inlet swirl angle. The air flow AA is “lifted” or guided by a pressure surface 13 to a desired direction.

Unfortunately, there are situations where there are significant inlet distortions to the presented air flow AA which produce a significant variation in the inlet flow angle, both circumferentially and radially. Such problems with variation in the inlet swirl angle are exacerbated when the inlet distortions and the aerofoils of the guide vanes are in the same frame of reference whether stationery or rotary. For example, such inlet distortions and inlet swirl angle variations can vary widely dependent upon the combustion system design of an engine incorporating a nozzle guide vane. In short, the angle of air flow AA presentation to the nozzle guide vane 12 alters the stagnation point 10 position and therefore the most appropriate distribution of apertures 11 in order to provide film cooling about aerofoil surfaces of that vane 12. The variation of inlet swirl around the annulus upon which the vanes are mounted will depend upon the ratio of combustion burners to nozzle guide vanes as well as burner settings at reduced power configurations. Variations in inlet angle to the nozzle guide vanes 12 will affect the leading edge stagnation point 10 and so the air flow from the apertures 11 in order to provide film cool about the vane.

As indicated above, FIG. Prior Art 1 illustrates a typical high pressure nozzle guide vane in schematic plan cross-section. It will be noted from FIG. 1 that the nozzle guide vanes 2 in the arrangement are set at an angle to the air flow A. If the number of aerofoils which create the nozzle guide vane arrangement is reduced and the chord CAX is fixed then a high lift design of nozzle guide vane is required.

In attached FIG. Prior Art 2 a possible schematic plan cross-section of a high lift nozzle guide vane 22 is illustrated. Thus, an air flow BB is again presented to the vane 22 such that the flow is diverted over the vane 22 as illustrated by the arrowheads. Apertures 24 are again provided within the vane 22 from which high pressure cooling air emanates in order to provide film cooling of the aerofoil surfaces of the vane 22. Of particular concern with respect to the present invention is that an aerofoil vane 22 as depicted in FIG. Prior Art 2 is susceptible to large displacements in leading edge stagnation point 20 position upon the vane 22 due to the relatively flat incident surface 21 upon which the air flow BB is presented. Variations in the inlet swirl angle of this flow BB will create movement of the stagnation point 20 up and down upon this surface 21. Variations in the leading edge stagnation point 20 render it more likely that the film cooling provided through high pressure cooling air flows from the apertures 24 will not follow the intended path upon the surface 21 or a suction surface 23 and so result in potentially unacceptably high operating temperatures or thermal cycling. Unsteady losses may also be increased due to the stagnation point movements leading to a loss of efficiency.

In accordance with the present invention there is provided a guide vane for an engine comprising a surface towards which an air flow is directed in use, the surface being curved with a leading edge bulge ridge to limit the proportion of surface in a flat presentation aspect to the air flow in use and so limit in use possible stagnation point positions for that flow upon the aerofoil.

Typically, the bulge ridge is located substantially centrally upon the surface of the vane.

Preferably, the bulge ridge is positioned off-centre. Possibly, the curves either side of the bulge ridge are differently shaped. Possibly, one curve is substantially flat whilst the other slope is curved.

Possibly, one slope upon the surface develops into an early suction side for the vane due to its shape and/or position relative to the bulge ridge.

Also in accordance with the present invention there is provided a nozzle guide vane assembly for a turbine engine incorporating a plurality of vanes as described above.

Further in accordance with the present invention there is provided an engine incorporating a nozzle guide vane assembly or nozzle guide vane as described above.

An embodiment of the present invention will be described by way of example only with reference to the accompanying drawings and in particular FIG. 2 schematically illustrating a plan cross-section of a vane in accordance with the present invention.

Referring to FIG. 3 which illustrates a vane 31 in accordance with the present invention. Thus, the vane 31 incorporates a pressure side 32 and a suction side 33 with an air flow C presented to the aerofoil. The aerofoil incorporates apertures 34 through which cooling air is released from supply channels 35. As can be seen from the flow stream lines marked with arrowheads the air flow C impinges or strikes upon the aerofoil such that there is a stagnation point 30 where the air flow C is substantially perpendicularly presented to the aerofoil. Air flows either side of the stagnation point 30 pass about and over the pressure side 32 and around a leading edge 36 onto the suction side 33.

At the stagnation point 30 as indicated previously, the forward air flow C is presented to the aerofoil and “rebounds” to create stagnateion through interaction with the forward flow. In such circumstances, it will be appreciated that the stagnation point 30 is a drag upon air flow. However, if the stagnation point 30 position is predictable and stable it will be appreciated by appropriate positioning of the apertures 32 adequate cooling by film overflow upon the pressure surface 32 and suction surface 33 can still be achieved. Previous vanes 12, 22 described respectively in FIG. Prior Art 1 and FIG. Prior Art 2 are not specifically configured in order to achieve a stable stagnation point position. In FIG. Prior Art 1 as indicated an acceptable stagnation position 10 is determined but only by providing a relatively low lift aerofoil design. With higher lift aerofoil designs as described with regard to vane 22 in FIG. Prior Art 2 the increased flat pressure surface 21 causes migration of the stagnation point 20 around the aerofoil such that specific distribution of the apertures in order to maintain cooling despite the effects of stagnation point 20 movement cannot be achieved as one position of the stagnation point requires one distribution of apertures whilst other positions will require different positions of the apertures.

In accordance with the present invention as depicted in FIG. 2 the part of aerofoil which within the pressure surface 32 is formed also incorporates a bulge ridge 37 from which extend slopes 32 a, 32 b. In such circumstances, the bulge ridge 37 provides a limited top surface which may be perpendicular to the presented air flow C. The slopes 32 a, 32 b either side of the bulge ridge 37 are curved as presented to the air flow C and so the air flow C is generally deflected. In such circumstances, the bulge ridge 37 and a limited margin either side of that ridge 37 constitute the potential range of positions for the stagnation point 30. In short, the position of the stagnation point 30 is well defined and stable such that there is a predictability which allows appropriate positioning of the apertures 34 in order to achieve film cooling as require. Normally, a deviation in the air flow C in accordance with the primary function of a nozzle guide vane is still achieved principally by the lower slope 32 b and more fundamentally by the suction side 33 which is substantially the same as the suction sides 14, 23 of previous vanes described with regard to FIG. Prior Art 1 and FIG. Prior Art 2.

It will be noted that in comparison with the previous vanes 12, 22 the vane 31 described in FIG. 2 has an increased cross-sectional area in coolant air supply channels 35 which allows an improved cooling air pressure to be presented through the apertures 34 for film cooling efficiency. Furthermore, the increased boxed shape of the vane 31 will improve mechanical strength and stiffness of the vane 31.

Returning to FIG. 1 illustrating a typical nozzle guide vane and turbine blade cooling arrangement 1 for a turbine engine. It will be noted that the vanes 2 are mounted such that they are angularly presented to the air flow A. This axial chord spacing CAX will typically be fixed as described previously such that if it is desired to reduce the number of vanes and aerofoils in a nozzle guide vane arrangement it is necessary for each vane to create more lift in the presented air flow A. As indicated previously, it is when attempting to achieve this greater lift that particular difficulties occur with regard to stagnation point predictability and stability. Nevertheless, in accordance with the present invention the bulge ridge 37 is defined as a forward part of the pressure surface 32 such that dotted line 38 constitutes the notional front edge of the nozzle guide vane surface presented to the air flow. Thus, the expected inlet flow angle will be towards that nominal front edge (broken line 38) and the slopes 32 a, 32 b will each be angularly curved in presentation to the flow. With variations in the inlet swirl angle for the reasons described previously, the actual presentation of the air flow C may swirl but nevertheless in accordance with the present invention the stagnation point 30 will remain substantially about the top of the bulge ridge 37.

Whilst endeavouring in the foregoing specification to draw attention to those features of the invention believed to be of particular importance it should be understood that the Applicant claims protection in respect of any patentable feature or combination of features hereinbefore referred to and/or shown in the drawings whether or not particular emphasis has been placed thereon. 

1. A guide vane for an engine comprising a pressure surface towards which an air flow is directed in use, the surface being curved with a leading edge bulge ridge to limit the proportion of pressure surface in a flat presentation aspect to the air flow in use and so limit in use possible stagnation point position for that flow upon the vane.
 2. A vane as claimed in claim 1 wherein the bulge ridge is located substantially centrally upon the surface of the vane.
 3. A vane as claimed in claim 1 wherein the bulge ridge is positioned off-centre upon the surface of the vane.
 4. A vane as claimed in claim 1 wherein slopes extend either side of the bulge ridge and the slopes either side of the bulge ridge are one of differently shaped and symmetrical.
 5. A vane as claimed in claim 4 wherein one slope is substantially flat whilst the other slope is trailingly curved.
 6. A vane as claimed in claim 1 wherein a portion of the surface develops into an early suction side for the vane due to one of its shape and position of that portion of the surface relative to the bulge ridge.
 7. A nozzle guide vane arrangement for a turbine engine incorporating a plurality of vanes as claimed in claim
 1. 8. An engine incorporating a nozzle guide vane arrangement as claimed in claim
 8. 